1. Field of the Invention
The present invention relates to an improved shroud assembly for high pressure stages of axial flow compressors and turbines such as are incorporated in gas turbine engines for aircraft.
"Radially", in the context of this specification, means a direction at right angles to the longitudinal axis of the engine, "upstream" means in the direction of the air intake of the engine, "downstream" means in the direction of the engine exhaust, and "circumferentially" refers to the locus traced by the end of a radius rotating about and at right angles to the longitudinal axis of the engine.
Axial flow compressor or turbine rotor blade stages operating at high gas temperatures in gas turbine engines are now being provided with specially designed shroud rings for the purpose of maintaining more nearly optimum clearances between the tips of the rotor blades and the shrouds over as wide a range of rotor speeds and temperatures as possible. The importance of this lies in that blade tip clearances or clearance gaps that are too large reduce the efficiency of the compressor or turbine whilst clearances which are too small may cause damage under some conditions due to interference between the blade tips and the shroud ring.
2. Description of the Prior Art
A known method of maintaining optimum blade tip clearances over a wide range of conditions involves matching the thermal response of the shroud ring and its supporting structure--in terms of increase or decrease of diameter with operating temperature--to the radial growth or shrinkage of the compressor or turbine rotor due to changing centrifugal forces and temperatures. In order to achieve this required matching, the shroud rings are composed of a number of segments, each describing a relatively short arc length circumferentially of the rotor stage.
Such shroud segments are individually connected to the supporting structure surrounding the shroud ring. For instance, the casing round the turbine blades is normally made up from a number of shroud segments each supported by adjacent nozzle guide vane support structures. An increase in the temperature of the gas stream causes thermal expansion of the guide vane support structures, thus causing the shrouds to move radially outwards. The tip clearance between the rotor blades and the shrouds is thereby increased, bringing about an associated drop in turbine efficiency.
However, in gas turbine engines a tip clearance gap has to exist in order that the rotor tips keep clear of the shrouds under various operating conditions. It is usual to adopt a compromise whereby the tip clearance is large enough to avoid contact between the rotor tips and the shrouds but is made as small as possible for maximum efficiency.
A problem that further arises in the design of shroud segments individually connected to a supporting structure is excessive sealing clearance between a shroud segment and its supporting structure. This excessive sealing clearance can arise because of manufacturing tolerances in the production of the shroud segments and the supporting structure, and because of differing thermal expansion or expansion rates between the two types of components as the operating temperatures change.
In the case of compressors, excessive sealing clearances cause decreased efficiency because they allow air on the high pressure side of the rotor to leak between the shroud segments and the supporting structure to the low pressure side of the rotor. In the case of turbines, excessive sealing clearances increase the consumption of the high pressure cooling air which is fed to the shroud segments and the adjacent components to cool them. This reduces the efficiency of the engine. Large sealing clearances also decrease the effectiveness of the cooling air in cooling the shroud segments by allowing cooling air to escape which would otherwise pass through small cooling air passages in the shroud segments.
An object of the present invention is to provide an improved shroud assembly in which the segmented shroud members are supported in such a manner that distortion of the nozzle guide vanes brought about by thermal or other means has a minimal effect on the clearances between shroud members and rotor tips.